1. Field of the Invention
The present invention relates to a combustion gas turbine and, specifically, it relates to a split ring disposed on the inner wall surface of a gas turbine casing.
2. Description of the Related Art
A turbine casing of a combustion gas turbine forms a hot gas path through which high temperature combustion gas passes. Therefore, a lining made of a heat resistant material (such as a thermal protection tile) is disposed on the inner wall surface in order to prevent the casing metal surface from directly contacting hot combustion gas. Usually, the thermal protection lining is composed of a plurality of split segments arranged on the inner surface of the turbine casing in a circumferential direction so that the segments form a ring. Therefore, the thermal protection lining of the turbine casing is often called xe2x80x9ca split ringxe2x80x9d. In order to avoid problems due to thermal expansion at a high temperature, the respective split segments are spaced apart from each other in a circumferential direction.
FIG. 1 shows a cross-section of a turbine casing taken along the center axis thereof which indicates the position of the split ring.
In FIG. 1, numeral 1 designates a turbine casing as a whole. The turbine casing 1 has a cylindrical form in which a plurality of annular casing segments 3 made of metal are joined to each other in the axial direction.
Each casing segment is provided with a thermal insulation ring 5 disposed inside the casing segment 3 and spaced apart from the inner surface of the casing segment 3. Stator blades 9 of the respective turbine stages are fixed to the thermal insulation ring 5 through a stator ring 7.
Further, a split ring 10 is attached to the inner surface of each thermal insulation ring 5 at the portion between the stator rings 7 in such a manner that the inner surface of the split ring 10 opposes the tips of the rotor blades 8 with a predetermined clearance therebetween.
The split ring 10 is, as explained before, composed of a plurality of split segments made of a heat resistant material and arranged in the circumferencial direction of the casing inner wall. The respective split segments are spaced apart, in the circumferential direction, at a predetermined distance in order to accommodate the thermal expansion of the split segments.
A split ring of this type is disclosed in, for example, Japanese Unexamined Patent Publication (Kokai) No. 2000-257447.
The split segment of the split ring in the ""447 publication is provided with an internal cooling air passage for cooling the split segment. Cooling air after cooling the split segment is injected from the outlet of the passage disposed on the end face of the split segment located on the downstream side thereof with respect to the direction of the rotation of the turbine rotor. The cooling air is injected from the above-noted outlet obliquely toward the end face of the adjacent split segment. Further, the comer between the end face located upstream side with respect to the direction of rotation of the rotor and the inner face of the split segment in the ""447 publication is cut off so that the cooling airxe2x80x94injected from the adjacent split segment flows along the inclined surface formed at the comer. Thus, the inclined surface between the end face and the inner face is cooled by the film of cooling air.
However, in the split ring composed of the split segments, heat load exerted on the corner of the split segment between the upstream end face and inner surface thereof is very high and, in some case, cooling by the cooling air film is not sufficient.
This problem will be explained with reference to FIG. 9.
FIG. 9 schematically illustrates a cross-section of the turbine casing perpendicular to its axis.
In FIG. 9, numeral 1 designates a turbine casing (more precisely, a thermal insulation ring), 11 designates split segments of the split ring 10. As explained before, the respective split segments 10 are arranged in the circumferential direction with relatively small clearance 13 therebetween. The rotor blades 8 rotate in the direction indicated by the arrow R with a small clearance between the inner face 11c of the split segments 11 and the tips of the rotor blades 8.
High temperature combustion gas flows through the casing 1 in the axial direction as a whole. However, when combustion gas passes through the rotor blades 8, a circumferential velocity component is given to combustion gas by the rotor blade rotation and combustion gas flows in the circumferential direction with a velocity substantially the same as the tip velocity of rotor blades in the clearance between the tips of the blades 8 and the split segment 11.
When this swirl flow of combustion gas passes the clearance 13 between the split segments 11, turbulence occurs in the swirl flow.
FIG. 10 schematically illustrates the behavior of the swirl flow FR of combustion gas when it passes the rotor blade 8. As shown in FIG. 10, when the swirl flow FR passes through the clearance 13 between the split segments 11, the swirl flow FR impinges on the lower portion (i.e., the portion near the corner between the end face and the inner face) of the upstream end faces lla of the split segment 11 before it flows into the clearance 13. Therefore, at the portion where swirl flow FR of combustion gas impinges on the upstream end face 11a, heat is transferred from combustion gas to the end face by an impingement heat transfer. This causes the heat transfer rate between the end face 11a and combustion gas flow FR to increase largely compared with the case where combustion gas flows along the inner face 11c of the split segments 11.
Due to this increase in the heat transfer rate, the lower portion of the upstream end face 11a (i.e., the portion near the corner between the upstream end face 11a and the inner face 11c) of the split segment 11 receives a large quantity of heat every time the rotor blade 8 passes the clearance 13. Therefore, the temperature of the corner portion of the upstream end faces 11a of the split segments 11 largely increases and, due to sharp increase in the local temperature, burning or cracking occurs at the corner portions of the split segments 11.
In the above-noted ""447 publication, since cooling air is injected and flows along the corner portion of the split segment, the temperature rise of the corner portion is suppressed to some extent. However, in the actual operation, since the flow of cooling air is disturbed by the impinging swirl flow of combustion gas, a cooling air film sufficient for cooling the corner portion is not formed and, thereby, cooling of the corner portion is insufficient even if the cooling air is supplied to the corner portion as disclosed by ""447 publication.
In view of the problems in the related art as set forth above, the objects of the present invention is to provide a split ring of a gas turbine casing capable of preventing the burning of the corner portion of the split segment by reducing the temperature rise caused by the impingement of the swirl flow of combustion gas.
The objects as set forth above is achieved by a split ring for a gas turbine casing, according to the present invention, comprising a plurality of split segments arranged on an inner wall of a gas turbine casing in a circumferential direction at predetermined intervals so that the split segments form a ring disposed between tips of turbine rotors and inner wall casing opposing the tips of the rotor blades, wherein each of the split segments includes two circumferential end faces which oppose the end faces of the adjacent split segments and an inner face substantially perpendicular to the end faces and opposing the tips of the rotors and a transition face formed between at least one of the end faces and the inner face and, wherein the surface of the transition face is formed in such a manner that the clearance between the tips of the rotor blades and the surface of the transition face increases from the inner face toward the end face.
According to the present invention, at least one of the end faces of the split segment is connected to the inner face by a transition face.
When the transition face is formed between the upstream end face and the inner face, the swirl flow of combustion gas flows along the transition face and does not impinge the end face. Therefore, an increase in the heat transfer rate on the end face does not occur.
When the transition face is formed between the downstream end face and the inner face, as the cross-section of the flow path of the swirl flow (i.e. the clearance between the tips of the rotor blades and the transition face) increases as it approaches the downstream end face. Therefore, the circumferential velocity of the swirl flow decreases near the downstream end face due to diversion of the flow passage. Thus, when the rotor blade passes the clearance between the split segments, though the swirl flow still impinges the upstream end face of the split segments, the velocity of the swirl flow when it impinges the end face is largely reduced and the increase in the heat rate due to impingement is suppressed.
As explained above, the transition face can be disposed either between the upstream end face and the inner face or between the downstream end face and the inner face. Further, the transition face can be disposed between inner face and both of the end faces.
The surface of the transient face can be any shape as long as the clearance between the rotor blade tip and the transition face increases from the end face toward the inner face. The transition face may be formed as a plane oblique to inner face and the end face. Further, the transition face may be formed as a cylindrical surface or a spherical surface.